Composite structures are used in a wide variety of applications due to their high strength-to-weight ratio, improved corrosion resistance, and other favorable properties. In aircraft construction, composites are used in increasing quantities to form the fuselage, wings, horizontal and vertical stabilizer, and other components. For example, the horizontal stabilizer of an aircraft may be formed of composite skin panels co-bonded to internal composite structures such as composite stiffeners or spars. The composite spars may extend from the root to the tip of the horizontal stabilizer and may generally taper in thickness along a spanwise direction to improve the stiffness characteristics of the horizontal stabilizer and reduce weight. Composite spars may also include localized increases in the composite thickness such as where the spar attaches to other structures or components.
Composite stiffeners or spars may be provided in a variety of cross-sectional shapes. For example, a composite spar or stiffener may be formed in an I-beam shape by bonding together the vertical webs of two composite C-channels in back-to-back arrangement. Each one of the C-channels may have horizontal flanges extending outwardly from the upper and lower sides of the web. Each horizontal flange may transition into the web at a radiused web-flange transition or stiffener radius. When the C-channels are joined back-to-back to form the I-beam shape, the back-to-back stiffener radii result in a lengthwise notch region along each one of the upper and lower sides of the I-beam. To improve the strength, stiffness, and durability of a composite structure, the notch regions may be filled with a radius filler formed of composite material.
Unfortunately, existing radius fillers suffer from several drawbacks that detract from their overall utility. For example, existing radius fillers may exhibit microcracking during manufacturing such as during cool-down from curing or during trimming operations. Microcracking may also occur during service under certain loading conditions and/or during thermal cycling. Microcracking in conventional radius fillers may be difficult to detect during routine non-destructive inspection due to the orientation of the microcracks which may be perpendicular to the composite skin. Furthermore, conventional radius fillers may exhibit a relatively low pull-off strength at the bond between the I-beam and the skin panel.
A further drawback associated with conventional radius fillers is the inability to vary the mechanical properties of the radius filler in multiple directions as may be desired in composite assemblies that are subjected to different loading conditions at different locations. For example, a homogenous radius filler formed of bundles of unidirectional fiber tows cannot be configured to have relatively high stiffness at one lengthwise location of the radius filler and have high strain capability (e.g., relatively low stiffness) at another lengthwise location.
As can be seen, there exists a need in the art for a radius filler that minimizes microcracking during manufacturing, in service, and during thermal cycling, and which provides favorable pull-off strength. In addition, there exists a need in the art for a radius filler that provides the capability to tailor the mechanical properties of the radius filler in one or more directions.